A gas turbine, or jet, engine may be generalized as having three overall sections: 1) a compressor which receives and compresses incoming gas such as air, 2) a combustion chamber wherein the compressed gas is mixed with fuel and burned to produce exhaust gas; and 3) one or more turbines which extract energy from the high-pressure, high-velocity exhaust gas exiting the combustion chamber.
The arrangement and configuration of these three components impacts many characteristics of the gas turbine engine, including overall engine length and weight, as well as the materials to construct the turbine engine. The overall length of the turbine engine may be shortened, saving on materials, weight and length, by the use of a reverse flow annular combustion chamber. This type of combustion chamber is so named because the mean direction of flow within the chamber is opposite the general direction of air flow through the engine as a whole. A transition liner assembly is fitted to the downstream portion of the annular combustion chamber and serves to redirect the flow of combustion gas into the turbine section resulting in a gas flow aligned with the general direction of overall flow through the engine.
FIG. 1 shows a functional schematic cut-away diagram of one conventional gas turbine engine 10, which includes an air compressing passage 14, a combustion chamber 30, and a turbine 20. More specifically, FIG. 1 shows a portion of the combustor-to-turbine transition assembly which is generally comprised of a transition liner assembly 51 mounted adjacent to a combustion chamber 30 so that the flow of combustion gasses, represented by arrows 68, is diverted from the combustion chamber 30 to the turbine 20 section of the engine. More specifically, in this particular engine, the direction of the combustion gas flow is reversed with respect to the orientation of the turbine engine 10 as the combustion gas is directed from the upstream section 55 of the transition liner 53 to the downstream section 57 of the transition liner 53 to the direction of flow through the turbine 20. To assemble the above components, the inner edge of the upstream section 55 of the transition liner assembly 51 may be generally slip fit, or otherwise fastened, over or adjacent to the outer edge 34 of the combustion chamber 30 downstream section 32. To improve the longevity and performance of the combustion chamber, the outer surface of the downstream section 32 of the combustion chamber 30 can be coated with a hardening material, which can be a chrome carbide metal spray.
The forces applied to an operational gas turbine engine, especially when affixed to the wing of an aircraft, can result in vibration between internal engine components. Vibration between the transition liner and the hardened edge of the combustion chamber 30 can result in damage or wear (otherwise called erosion or fretting) of the edge of the transition liner.
The prior art depicted in FIG. 1B shows erosion damage in the form of a ledge 59 that has been worn into the upstream section 55 of an eroded transition liner 53. As shown, the erosion-induced ledge 59 has degraded the connection between the upstream section 55 of transition liner 53 and the outer downstream section 32 of the combustion chamber 30, such that some combustion gasses 68 may escape, represented by arrows 65. Akin to a blown head gasket in a conventional automobile engine, the escape of gasses in an undesired location results in degraded engine performance and may shorten the operational life expectancy of the turbine engine or downtime and cost related to engine maintenance. However, it should be appreciated that despite this drawback, conventional gas turbine engines generally are operationally safe and reliable.
Upon a loss of gas turbine engine performance or as required by routine turbine engine maintenance, disassembly of the turbine engine occurs and the transition liner assembly may be inspected. When an upstream annular section of the transition liner is identified as being damaged by erosion (for example by visual inspection), the entire transition liner is removed and replaced. The replacement transition liner is subject to the same erosion damage and life expectancy drawbacks as the old eroded liner. Thus, eroded transition liners may cause maintenance expense and engine downtime.
Hence, there is a need in gas turbine engines for a combustor-to-turbine transition liner assembly, and a related repair method, that overcomes one or more of the drawbacks identified above. The present invention satisfies one or more of these needs.